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NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA/LANGLEY RC-08(B)3 AIRFOIL (rc08b3-il)
Reynolds number: 100,000
Max Cl/Cd: 45.51 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rc08b3-il-100000.txt
Download as CSV file: xf-rc08b3-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/LANGLEY RC-08(B)3 AIRFOIL                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.4681   0.10459   0.09980  -0.0101   1.0000   0.0976
  -9.500  -0.4885   0.10112   0.09640  -0.0148   1.0000   0.0989
  -9.250  -0.5094   0.09708   0.09243  -0.0200   1.0000   0.0992
  -9.000  -0.4672   0.09204   0.08733  -0.0121   1.0000   0.1042
  -8.750  -0.4676   0.08820   0.08351  -0.0129   1.0000   0.1079
  -8.500  -0.4806   0.08394   0.07931  -0.0162   1.0000   0.1115
  -8.250  -0.6062   0.08775   0.08293  -0.0166   1.0000   0.1009
  -8.000  -0.5902   0.08458   0.07978  -0.0134   1.0000   0.1047
  -7.750  -0.5902   0.08067   0.07586  -0.0156   1.0000   0.1091
  -7.500  -0.6160   0.07769   0.07247  -0.0226   1.0000   0.1138
  -7.250  -0.5920   0.07196   0.06704  -0.0196   1.0000   0.1170
  -7.000  -0.5833   0.06871   0.06376  -0.0193   1.0000   0.1226
  -6.750  -0.5851   0.06459   0.05944  -0.0209   1.0000   0.1301
  -6.500  -0.5718   0.06147   0.05636  -0.0195   1.0000   0.1359
  -6.250  -0.5676   0.05795   0.05267  -0.0194   1.0000   0.1457
  -6.000  -0.5621   0.05531   0.04980  -0.0185   1.0000   0.1585
  -5.750  -0.5538   0.05275   0.04715  -0.0168   1.0000   0.1728
  -5.500  -0.5439   0.04991   0.04434  -0.0148   1.0000   0.1882
  -5.250  -0.5341   0.04729   0.04176  -0.0126   1.0000   0.2043
  -5.000  -0.5242   0.04488   0.03935  -0.0103   1.0000   0.2215
  -4.500  -0.4764   0.03264   0.02455  -0.0062   1.0000   0.0836
  -4.250  -0.4585   0.02947   0.02114  -0.0043   1.0000   0.0780
  -4.000  -0.4371   0.02763   0.01847  -0.0014   1.0000   0.0717
  -3.750  -0.4160   0.02582   0.01637   0.0001   1.0000   0.0713
  -3.500  -0.3944   0.02347   0.01387   0.0011   1.0000   0.0740
  -3.250  -0.3716   0.02208   0.01237   0.0022   1.0000   0.0766
  -3.000  -0.3474   0.02074   0.01089   0.0033   1.0000   0.0783
  -2.750  -0.3232   0.01960   0.00965   0.0044   1.0000   0.0817
  -2.500  -0.2996   0.01846   0.00846   0.0054   1.0000   0.0872
  -2.250  -0.0850   0.01313   0.00640  -0.0271   1.0000   1.0000
  -2.000  -0.0717   0.01308   0.00618  -0.0248   1.0000   1.0000
  -1.750  -0.0604   0.01308   0.00604  -0.0221   1.0000   1.0000
  -1.500  -0.0503   0.01311   0.00595  -0.0192   1.0000   1.0000
  -1.250  -0.0411   0.01318   0.00591  -0.0161   1.0000   1.0000
  -1.000  -0.0322   0.01328   0.00590  -0.0130   1.0000   1.0000
  -0.750  -0.0230   0.01341   0.00593  -0.0100   1.0000   1.0000
  -0.500  -0.0133   0.01355   0.00599  -0.0071   1.0000   1.0000
  -0.250  -0.0030   0.01372   0.00609  -0.0044   1.0000   1.0000
   0.000   0.0080   0.01392   0.00622  -0.0019   1.0000   1.0000
   0.250   0.0198   0.01414   0.00637   0.0005   1.0000   1.0000
   0.500   0.0320   0.01439   0.00657   0.0026   1.0000   1.0000
   0.750   0.0450   0.01467   0.00681   0.0046   1.0000   1.0000
   1.000   0.0588   0.01499   0.00710   0.0063   1.0000   1.0000
   1.250   0.1116   0.01539   0.00751   0.0004   0.9901   1.0000
   1.500   0.1662   0.01576   0.00791  -0.0057   0.9790   1.0000
   1.750   0.2232   0.01604   0.00827  -0.0122   0.9674   1.0000
   2.000   0.2845   0.01618   0.00852  -0.0192   0.9552   1.0000
   2.250   0.3462   0.01615   0.00865  -0.0260   0.9418   1.0000
   2.500   0.4101   0.01590   0.00862  -0.0328   0.9271   1.0000
   2.750   0.4563   0.01558   0.00849  -0.0356   0.9057   1.0000
   3.000   0.5005   0.01507   0.00818  -0.0373   0.8842   1.0000
   3.250   0.5283   0.01463   0.00786  -0.0355   0.8563   1.0000
   3.500   0.5504   0.01415   0.00747  -0.0323   0.8230   1.0000
   3.750   0.5698   0.01374   0.00709  -0.0286   0.7823   1.0000
   4.000   0.5885   0.01350   0.00687  -0.0251   0.7302   1.0000
   4.250   0.6076   0.01343   0.00668  -0.0219   0.6614   1.0000
   4.500   0.6249   0.01373   0.00658  -0.0186   0.5534   1.0000
   4.750   0.6373   0.01488   0.00681  -0.0151   0.3815   1.0000
   5.000   0.6448   0.01729   0.00783  -0.0120   0.1850   1.0000
   5.250   0.6599   0.01900   0.00907  -0.0099   0.1334   1.0000
   5.500   0.6787   0.02032   0.01029  -0.0081   0.1123   1.0000
   5.750   0.6989   0.02163   0.01152  -0.0066   0.0987   1.0000
   6.000   0.7210   0.02339   0.01311  -0.0055   0.0909   1.0000
   6.250   0.7444   0.02471   0.01463  -0.0043   0.0843   1.0000
   6.500   0.7674   0.02683   0.01667  -0.0036   0.0779   1.0000
   6.750   0.7905   0.02859   0.01881  -0.0022   0.0754   1.0000
   7.000   0.8123   0.03080   0.02142  -0.0007   0.0734   1.0000
   7.250   0.8319   0.03285   0.02377   0.0009   0.0701   1.0000
   7.500   0.8508   0.03517   0.02620   0.0021   0.0671   1.0000
   7.750   0.8669   0.03829   0.02976   0.0041   0.0670   1.0000
   8.000   0.8737   0.04259   0.03499   0.0076   0.0699   1.0000
   8.250   0.8785   0.04730   0.04028   0.0104   0.0728   1.0000
   8.500   0.8813   0.05175   0.04514   0.0128   0.0744   1.0000
   8.750   0.8837   0.05636   0.05003   0.0147   0.0761   1.0000
   9.500   0.7393   0.09258   0.08753   0.0002   0.1668   1.0000
   9.750   0.7882   0.09379   0.08882   0.0098   0.1601   1.0000
  10.000   0.7280   0.10215   0.09701  -0.0034   0.1563   1.0000
  10.250   0.7274   0.10620   0.10105  -0.0044   0.1504   1.0000
  10.500   0.7424   0.11009   0.10496  -0.0017   0.1456   1.0000
  10.750   0.7157   0.11564   0.11039  -0.0086   0.1400   1.0000
  11.000   0.5922   0.11161   0.10660   0.0041   0.1482   1.0000
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